Compressor rotor with coated blades

ABSTRACT

A compressor rotor for a gas turbine engine has blades circumferentially distributed around and extending a span length from a central hub. The blades include alternating first and second blades having airfoils with corresponding geometric profiles. The airfoil of the first blade has a coating varying in thickness relative to the second blade to provide natural vibration frequencies different between the first and the second blades.

TECHNICAL FIELD

The application relates generally to rotating airfoils for gas turbineengines, and more particularly to mistuned compressor rotors.

BACKGROUND

Aerodynamic and/or vibrational instabilities, such as but not limited toflutter, can occurs in a gas turbine engine when two or more adjacentblades of a rotor of the engine, such as the fan, vibrate at a frequencyclose to their natural frequency and the interaction between adjacentblades maintains and/or strengthens such vibration. Other types ofaerodynamic instability, such as resonant response, may also occur andare undesirable. Prolonged operation of a rotor undergoing suchinstabilities can cause airfoil stress loads to exceed acceptable levelsfor operation.

Various attempts have been made to mistune adjacent blades of suchrotors so as to separate their natural frequencies and reduce thelikelihood of undesirable instabilities. Continuous improvement isnevertheless sought.

SUMMARY

There is accordingly provided a compressor rotor for a gas turbineengine, the rotor comprising blades circumferentially distributed aroundand extending a span length from a central hub, the blades includingalternating first and second blades having airfoils with a leading edge,a trailing edge, a root, a tip and a mid-span region midway between theroot and the tip along the span, the airfoils of the first and secondblades having corresponding geometric profiles, the airfoil of the firstblades having a coating defining a first coating structure, the coatingbeing provided on at least a portion of the first blade adjacent theroot and having a root coating thickness, the mid-span region of thefirst blade having a mid-span thickness, the coating being provided on aportion adjacent the tip of the first blade and having a tip coatingthickness, the root coating thickness being greater than at least one ofthe tip coating thickness and a coating thickness of the airfoil of thefirst blade at the mid-span region , the first coating structure of thefirst blade selected to provide the first blade with a first naturalvibration frequency different from a second natural vibration frequencyof the second blade.

There is also provided a method of manufacturing a compressor rotor of agas turbine engine, the rotor having a plurality of bladescircumferentially distributed around and extending a span length from acentral hub, the method comprising the steps of: providing first andsecond blades respectively having first and second airfoils withcorresponding geometric profiles, a leading edge, a trailing edge, aroot, a tip, and a mid-span region midway between the root and the tipalong the span; and applying a coating on an outer surface of the firstairfoil to form a first coating structure, including applying thecoating on a portion of the first airfoil adjacent the root to define aroot coating thickness and applying the coating on a portion adjacentthe tip to define a tip coating thickness, the root coating thicknessbeing greater than at least one of the tip coating thickness and acoating thickness of the airfoil of the first blade at the mid-spanregion, wherein the first coating structure of the first blade isselected to provide a first natural vibration frequency different from asecond natural vibration frequency of the second blade.

There is also provided a compressor rotor for a gas turbine engine, themistuned compressor rotor comprising blades circumferentiallydistributed around and extending a span length from a central hub, theblades including alternating first and second blades havingcorresponding geometric profiles, the first blade having airfoil with acoating thereon within one or more portions thereof and defining a firstcoating structure, the one or more portions of the airfoil including aradially inner portion of the airfoil adjacent a blade root of the firstblade and having a root coating thickness and a radially outer portionof the airfoil adjacent a blade tip of the first blade and having a tipcoating thickness, the root coating thickness being greater than atleast one of the tip coating thickness and a mid-span coating thicknessat a mid-span region of the airfoil of the first blade, the firstcoating structure of the first blade selected to provide a first naturalvibration frequency different from a second natural vibration frequencyof the second blade.

There is further provided a rotor blade for a compressor rotor of a gasturbine engine, the compressor rotor having alternating blades havingcorresponding geometric profiles but different coating structuresselected to provide different vibration frequencies, the rotor bladecomprising an airfoil having a blade root, a blade tip, a coating on oneor more portions of the airfoil and defining a first coating structure,the one or more portions of the airfoil including a radially innerportion adjacent the blade root and having a root coating thickness anda radially outer portion adjacent the blade tip and having a tip coatingthickness, the root coating thickness being greater than at least one ofthe tip coating thickness and a mid-span coating thickness defined at amid-span region midway between the blade root and the blade tip.

BRIEF DESCRIPTION OF THE DRAWINGS

Reference is now made to the accompanying figures in which:

FIG. 1 is a schematic cross-sectional view of a gas turbine engine;

FIG. 2 is a perspective view of a fan rotor of the gas turbine engineshown in FIG. 1;

FIG. 3 is a side elevational view a first fan blade of the fan rotor ofFIG. 2;

FIG. 4A is a side elevational view the first fan blade of the rotor ofFIG. 2, according to an embodiment of the present disclosure;

FIG. 4B is a cross-sectional views of the first fan blade of FIG. 4Aalong axis A-A;

FIG. 4C is a cross-sectional views of the first fan blade of FIG. 4Aalong axis B-B;

FIG. 5A is a side elevational view the first fan blade of the rotor ofFIG. 2, according to an embodiment of the present disclosure;

FIG. 5B is a cross-sectional views of the first fan blade of FIG. 5Aalong axis A-A;

FIG. 5C is a cross-sectional views of the first fan blade of FIG. 5Aalong axis B-B; and

FIG. 5D is a cross-sectional view of the first fan blade of FIG. 5Aalong axis C-C.

DETAILED DESCRIPTION

FIG. 1 illustrates a gas turbine engine 10 of a type preferably providedfor use in subsonic flight, generally comprising in serial flowcommunication a fan 12 through which ambient air is propelled, acompressor section 14 for pressurizing the air, a combustor 16 in whichthe compressed air is mixed with fuel and ignited for generating anannular stream of hot combustion gases, and a turbine section 18 forextracting energy from the combustion gases. Engine 10 also comprises anacelle for containing various components of engine 10. Although theexample below is described as applied to a fan of a turbofan engine, itwill be understood the present teachings may be applied to any suitablegas turbine compressor rotor.

A compressor rotor assembly for the gas turbine engine 10 is provided.The rotor assembly as described herein can be the fan 12, or anothercompressor rotor of the compressor section 14, for example. In oneparticular embodiment, the fan 12 may be an Integrally Bladed Fan (IBF).

Referring to FIG. 2, the fan 12 includes a central hub 22, which in userotates about an axis of rotation 21, and a circumferential row of fanblades 24 that are circumferentially distributed and which project atotal span length L from hub 22 in a span-wise direction (which may besubstantially radially). The axis of rotation 21 of the fan 12 may becoaxial with the main engine axis 11 of the engine 10 as shown inFIG. 1. The fan 12 may be either a bladed rotor, wherein the fan blades24 are separately formed and fixed in place on the hub 22, or the fan 12may be an integrally bladed rotor (IBR), wherein the fan blades 24 areintegrally formed with the hub 22. Each circumferentially adjacent pairof fan blades 24 defines an inter-blade passages 26 there-between forthe working fluid.

In one particular embodiment, the compressor rotor and/or fan 12 asdescribed herein may be a “mistuned” rotor, in which the blades of therotor are not uniform about the circumference of the rotor andselected/designed such as to ensure a frequency separation between thedifferent blade types.

The circumferential row of fan blades 24 of fan 12 thus includes, in atleast one particular embodiment, two or more different types of fanblades 24, in the sense that a plurality of sets of blades are provided,each set having airfoils, which will be described in more details belowand illustrated in further figures. It is to be understood, however,that each of these sets of fan blades 24 may include more than twodifferent blade types, and need not comprise pairs, or even numbers, ofblade types. For example, each set of fan blades may include three ormore fan blades which differ from each other (e.g. a circumferentialdistribution of the fan blades may include, in circumferentiallysuccessive order, blade types: A, B, C, A, B, C; or A, B, C, D, A, B, C,D, etc., wherein each of the capitalized letters represent differenttypes of blades as described above).

The embodiment described below includes, for the sake of simplicity ofexplanation, a fan 12 having circumferentially alternating sets of fanblades each composed of only two different blade types, namely blades 28(e.g. blade type “A”) and 30 (e.g. blade type “B”). This constitutes,accordingly, a circumferential distribution of fan blades in thisexample which follows a circumferential sequence of blade types A, B, A,B, etc.

The blades 28 and 30 are geometrically similar and their respectiveairfoils have corresponding geometric profiles. However, blades 28 and30 have properties which are different from one another, notably theirdensity and/or stiffness distribution, as will be described and shown infurther details below. The different properties of the first and secondfan blades 28, 30 provide a natural vibrational frequency separationbetween the adjacent blades 28 and 30, which may be sufficient to reduceor impede unwanted vibration resonance between the blades (such asbending or torsion vibration , also known as heave and pitch).Regardless of the exact amount of frequency separation, the first andsecond fan blades 28 and 30 are therefore said to be intentionally“mistuned” relative to each other, in order to reduce the occurrenceand/or delay the onset, of flow-induced resonance. It is contemplatedthat the vibration referred to in the present disclosure is a naturalvibration.

Referring to FIG. 3, the fan blade 28 (blade type “A”) of the fan 12, isshown. The fan blade 28 has an airfoil 32 with leading edge 34, trailingedge 36, and tip 38, as well as a root 42 which can have a platform anda blade fixing or dovetail for engaging a fan hub (not shown). The blade38 also has a mid-span 44 midway between the root 42 and the tip 38along the span length. A mid-span region 45 of the airfoil is definedabout the mid-span 44. Fan blade 30 (blade type “B”) is not shown buthas an airfoil with a corresponding geometric profile and has thereforea similar geometry.

Flow-induced resonance refers to a situation where, during operation,adjacent vibrating blades transfer energy back and forth through the airmedium, which energy continually maintains and/or strengthens theblades' natural vibration mode. Fan blades have a number of oscillationpatterns (or vibration modes), any of which, if it gets excited and goesinto resonance, can result in flow induced resonance issues. In bendingvibration mode there is translational motion, into and out of thedirection of blade rotation, at the blade tip. In torsion vibrationmode, the blade is deformed by rotation along a longitudinal axisrelative to the span length.

The airfoils of the fan blades 28 and 30 comprise a substrate. Thesubstrate can be a metallic substrate selected from one of a titanium,platinum aluminum, nickel and/or iron base substrate. It is understoodthat the metallic substrate can comprise other materials, such asadditives or impurities. The substrate can also be a composite substrate(such as a fiber reinforced plastic) or a non-metallic substrate such asa polymer (polyether ketone or nylon for example).

According to an embodiment of the present disclosure, the airfoil of fanblade 28 has a coating on at least a portion of an outer surface 46 ofthe substrate forming the airfoil 32. The coating increases thestiffness and/or density of the fan blade 28 in the coated regions. Thecoating is located on at least a portion of the airfoil 32 adjacent theroot 48 (or root portion) and a portion of the airfoil adjacent the tip50 (or tip portion). In an embodiment, the root and tip portions includethe root and the tip, respectively. However the root and tip portionscan also be proximate the root and the tip without including therespective root and tip. As illustrated in the embodiment shown of FIG.3, the coating covers the entire outer surface 46 of airfoil 32.However, in alternate embodiments, the coating can also be located onlyon portions of the airfoil 32 outer surface 46, namely the portionsincluding at least radially outer portions 48 (proximate the tip 38) andradially inner portions 50 (proximate the hub 42).

The coating is composed of a material that provides increased stiffnessand/or density relative to that of the underlying substrate. It iscontemplated that the skilled person will choose a suitable type ofcoating (e.g. metal, nanocrystalline metal, carbon nanotube, composite,ceramic, etc.) as to provide greater stiffness and/or density. In aparticular embodiment, the coating is a nanocrystalline metal coating(i.e. a “nano coating”). In one embodiment, the nano coating can becomposed of a material different to that of the airfoil substrate. Thenano coating can provide for improved structural properties (stiffnessand density, for example) and for improved fatigue endurance of theairfoil. The nano coating metal grain size may range between about 2 nmand 5000 nm. The nano coating may be a nickel (Ni), copper (Cu),cobalt-phosphorous (CoP) or another suitable metal or metal alloy, suchas Co, Cr, Fe, Mo, Ti, W, or Zr. The nanocrystalline metal coating maybe composed of a pure or single metal, such as Ni or Co for example. Ina particular embodiment, the nano coating is nickel nano coating. Thecoating can also comprise a non-metallic coating or a composite coatingthat can increase stiffness and/or density relative to that of theunderlying substrate.

The coating applied on the airfoil defines a blade thickness. On theradially inner portion 50 adjacent the root 42, the coating defines aroot coating thickness T_(r) (hereafter referred to as root thicknessT_(r)), and on the radially outer portion 48 adjacent the tip 38, thecoating defines a tip coating thickness T_(t) (hereafter referred to astip thickness T_(t)). The coating can also be present on the mid-spanregion 45, as defined herein. In such scenario, the coating defines acoating thickness of the first blade at the mid-span region T_(ms)(hereafter referred to as the “mid-span thickness” T_(ms)). The coatinghas a specific coating structure on the outer surface 32. The rootthickness T_(r), tip thickness T_(t) and mid-span thickness T_(ms) aredefined as the distance between the outer surface 46 forming thesubstrate of the airfoil 32 and an outer surface of the coating. Inother word, the different thicknesses described herein refer to thethicknesses of the coating layer applied on the substrate on thecorresponding location. It is understood that where no coating isapplied the thickness will be zero. As used herein, the term “coatingstructure” refers to the arrangement and distribution of the coating onthe airfoil, i.e. the location, the shape and the thickness of thecoating layer.

In one embodiment, the root thickness T_(r) of blade 28 is greater thanone or both of the tip thickness T_(t) and the mid-span thicknessT_(ms). For example, the root thickness T_(r) can be greater than thetip thickness T_(t) and the mid-span thickness T_(ms), or the rootthickness T_(r) can be greater than the mid-span thickness T_(ms) whilebeing equal to the tip thickness T_(t). In certain embodiments, the rootthickness T_(r) is not lower than any other thickness over the outersurface of the airfoil. However, in alternate embodiments, when bendingfrequency may wish to be adjusted for example, mass is added to theblade tip only so that T_(r) is zero or T_(r) is less than T_(t).

The root thickness T_(r), tip thickness T_(t) and mid-span thicknessT_(ms) can be the same over an entire circumferential outer surface ofthe root portion, tip portion and mid-span region, respectively.However, the root thickness T_(r), tip thickness T_(t) and mid-spanthickness T_(ms) can also be different at the trailing edge and leadingedge than on the side airfoil faces (i.e. the portions of the airfoil onthe suction side and the pressure side thereof extending from theleading edge to the trailing edge on both sides of the airfoil) of theroot portion, tip portion and mid-span region, respectively.

The coating on the airfoil increases the stiffness and/or density of theblade on the coated region. The stiffness distribution and/or densitydistribution of the blade 28 is therefore different from that of thesame airfoil without the coating. Such difference of stiffness and/ordensity distribution causes a change in a natural vibration frequency ofthe fan blade 28 relative to the corresponding natural vibrationfrequency of the airfoil without coating. The vibration frequency can beone of the bending mode vibration frequency and the torsion modevibration frequency.

Blade 30 can also comprises a coating, as described herein. However, thecoating on blade 30 is present, if at all, in lower amount than on blade28 such that the change in stiffness and/or density distribution, andconsequently in natural vibration frequency, is different than for blade28. Therefore, if blade 30 has a coating, the coating structure of blade30 is different from the coating structure of blade 30. That is, atleast one of the location, the shape, and/or the thickness of thecoating layer may be different. Blade 30 can also be free from coatingso that the natural vibration frequency is that of the airfoil. As aresult, blade 28 has a first natural vibration frequency and blade 30has a second natural vibration frequency that is different from thefirst natural vibration frequency. There is therefore a frequencyseparation between blade 28 and blade 30, without a significant changeis the geometry of blades 28, 30. The frequency separation can be usedto dampen potential flutter or resonance between blades 28 and 30 and toprevent potential damages due to such vibration instabilities.

Considering that the airfoils of blades 28 and 30 have correspondinggeometric shapes/profiles, the different coating thicknesses (i.e. atleast T_(r), T_(t), T_(ms)) lead to small differences between theshapes/profiles of the blades 28, 30. However, the coating applied mayonly cause slight changes in the blade geometry, and massive changes maybe avoided. It is understood that the portions 48 and 50, and of themid-span region 45, illustrated by the dashed lines in FIG. 3, can be ofany suitable size and/or shape, which can be determined by the skilledpractitioner depending on at least the vibration mode considered, thenature of the airfoil substrate, and the nature of the coating.

Referring to FIGS. 4A-4C, a blade 128 in accordance with a particular isshown, where elements similar to that of the blade 28 are referred tousing the same reference numeral and will not be described furtherherein. In the particular embodiment of FIG. 4A, as illustrated by thecross sectional views taken along the axes A-A and B-B (FIGS. 4B and 4C,respectively), the root thickness T_(r) of blade 128 is greater than thetip thickness T₁ of the blade 128. In other words a thicker layer ofcoating is present toward the root (over a circumferential outer surfaceof root portion 50) than toward the tip (over a circumferential outersurface tip portion 48). It is understood that the coating on blade 130(not shown, similar to blade 30) can be located similarly than on blade128 (i.e. following a similar pattern, i.e. location and shape), butstill in a lower amount so that the coating structure of blade 128 isdifferent from the coating structure of blade 130. The coating can alsobe located uniformly over the whole airfoil of the blade 130 or can belocated following any other suitable pattern. Further, blade 130 can befree of coating. In any scenario, the bending vibration response ofblade 128 is changed and a vibrational frequency separation is createdbetween blade 128 and blade 130.

Referring to FIGS. 5A-5D, a blade 228 in accordance with anotherembodiment is shown, where elements similar to that of the blade 28 arereferred to using the same reference numeral and will not be describedfurther herein. In the particular embodiment of FIG. 5A, as illustratedby the cross sectional views taken along the axes A-A, B-B, and C-C(Figs. B to 5D, respectively), the root thickness T_(r) and the tipthickness T_(t) on the trailing and the leading edges of the blade 228is greater than the mid-span thickness T_(ms) on the trailing and theleading edges of the blade 228. In other words a thicker layer ofcoating is present on the leading edge and the trailing edge toward theroot (along the height of the root portion 50) and toward the tip (alongthe height of the tip portion 48) and a thinner layer of coating ispresent on the leading edge and trailing edge at the mid-span region 45.It is understood that the coating on blade 230 (not shown, similar toblade 30) can be located similarly than on blade 228 (i.e. following asimilar pattern, i.e. location and shape), but still in a lower amount,so that the coating structure of blade 228 is different from the coatingstructure of blade 230. The coating can also be located uniformly overthe whole airfoil of the blade 230 or can be located following any othersuitable pattern. Further, blade 230 can be free of coating. In anyscenario, the torsion vibration response of blade 228 is changed and avibrational frequency separation is created between blade 228 and blade230.

A method of manufacturing a compressor rotor assembly for a gas turbineengine is also provided. The compressor rotor assembly is as definedherein and comprises a plurality of blades each having an airfoil. Theairfoils are circumferentially distributed around and extend a spanlength from a central hub. The blades include successively alternatingtype “A” blade and type “B” blade. Type “A” blade and type “B” blade areprovided with corresponding geometric profiles, as defined herein.

Coating as defined herein is applied on at least an outer surface of theairfoil of type “A” blade, as to change the vibrational response(frequency) of at least type “A” blade. The coating is applied on atleast portions a portion adjacent the root and a portion adjacent thetip. Applying the coating comprises forming a coating layer on the outersurface of the airfoil of type “A” blade, thereby defining at least aroot thickness and a tip thickness delimited between an outer surface ofthe airfoil and an outer surface of the coating layer. In oneembodiment, the coating layer on the portion adjacent the root isthicker than the coating layer in one or both of the portion adjacentthe tip and the mid-span region, so that the root thickness is greaterthan one or both of the tip thickness and mid-span thickness. Thecoating is therefore applied to form a coating structure that induces achange in the vibrational response of the blade, relative to an airfoilwithout coating or with less coating.

In one embodiment, the coating is applied in greater amount (i.e. athicker coating layer) on the portion adjacent the root than on theportion adjacent the tip. The root thickness is therefore greater thanthe tip thickness. In another embodiment, the coating is applied ingreater amount (i.e. a thicker coating layer) on the leading andtrailing edges of the portions adjacent the root and the tip than on theleading and trailing edges of the mid-span region.

In one embodiment, the coating is applied on type “A” blade only, andtype “B” blade is kept free of coating. Therefore, the stiffness and/ordensity distribution of type “A” blade and type “B” blade, as well astheir respective vibration responses (frequencies), are different. Inanother embodiment the coating can also be applied on type “B” blade,but in a lower amount so that a coating structure of type “B” blade isdifferent from the coating structure of type “A” blade and the stiffnessand/or density distribution of type “A” blade and type “B” blade, aswell as their respective vibration responses (frequencies), aredifferent. The coating can be applied on the entire outer surface of theairfoil of type “B” blade. The coating can also be applied followingtype “A” blade coating pattern or any other suitable coating pattern,depending on at least the vibration mode considered, the nature of theairfoil substrate, and the nature of the coating.

The coating may be applied, according to the present method, through aplating process in a bath, such as to apply the fine-grained (i.e.nano-scale) metallic coating to the component or article to be coated.However, it is understood that any suitable plating or other coatingprocess, can be used. The coating can for example be applied using themethod described in U.S. Pat. No. 8,871,297, which is incorporatedherein by reference.

The above description is meant to be exemplary only, and one skilled inthe art will recognize that changes may be made to the embodimentsdescribed without departing from the scope of the invention disclosed.Any suitable base metals, polymers or other materials may be used as thesubstrate material, and any suitable metal and/or metal combinations maybe selected for the coating. Any suitable manner of applying the coatinglayer may be employed. Still other modifications which fall within thescope of the present invention will be apparent to those skilled in theart, in light of a review of this disclosure, and such modifications areintended to fall within the appended claims.

1. A compressor rotor for a gas turbine engine, the rotor comprisingblades circumferentially distributed around and extending a span lengthfrom a central hub, the blades including alternating first and secondblades having airfoils with a leading edge, a trailing edge, a root, atip and a mid-span region midway between the root and the tip along thespan, the airfoils of the first and second blades having correspondinggeometric profiles, the airfoil of the first blades having a coatingdefining a first coating structure, the coating being provided on atleast a portion of the first blade adjacent the root and having a rootcoating thickness, the mid-span region of the first blade having amid-span thickness, the coating being provided on a portion adjacent thetip of the first blade and having a tip coating thickness, the rootcoating thickness being greater than at least one of the tip coatingthickness and a coating thickness of the airfoil of the first blade atthe mid-span region , the first coating structure of the first bladeselected to provide the first blade with a first natural vibrationfrequency different from a second natural vibration frequency of thesecond blade.
 2. The compressor rotor as defined in claim 1, wherein theairfoil of the second blade is free of coating.
 3. The compressor rotoras defined in claim 1, wherein the airfoil of the second blade has acoating defining a second coating structure, the coating being providedon at least a portion adjacent the root of the second blade and having aroot coating thickness, the coating being provided on a portion adjacentthe tip of the second blade and having a tip coating thickness, the rootcoating thickness being greater than at least the tip coating thicknessand a coating thickness of the airfoil of the second blade at themid-span region, the second coating structure of the second bladeleading to the second natural vibration frequency that is different fromthe first natural vibration frequency.
 4. The compressor rotor asdefined in claim 1, wherein the root coating thickness of the firstblade is greater than the tip coating thickness of the first blade. 5.The compressor rotor as defined in claim 1, wherein the root coatingthickness and the tip coating thickness on the trailing and the leadingedges of the first blade is greater than the coating thickness of theairfoil at the mid-span region on the trailing and the leading edges ofthe first blade.
 6. The compressor rotor as defined in claim 1, whereinthe first and the second airfoil comprise a titanium based substrate orplatinum based substrate.
 7. The compressor rotor as defined in claim 1,wherein the coating is a nano-coating.
 8. The compressor rotor asdefined in claim 7, wherein the nano-coating is a nickel nano-coating.9. The compressor rotor as defined in claim 7, wherein the rotor is afan.
 10. A method of manufacturing a compressor rotor of a gas turbineengine, the rotor having a plurality of blades circumferentiallydistributed around and extending a span length from a central hub, themethod comprising the steps of: providing first and second bladesrespectively having first and second airfoils with correspondinggeometric profiles, a leading edge, a trailing edge, a root, a tip, anda mid-span region midway between the root and the tip along the span;and applying a coating on an outer surface of the first airfoil to forma first coating structure, including applying the coating on a portionof the first airfoil adjacent the root so that the first blade has aroot coating thickness and applying the coating on a portion adjacentthe tip so that the first blade has a tip coating thickness, the rootcoating thickness being greater than at least one of the tip coatingthickness and a coating thickness of the airfoil of the first blade atthe mid-span region, wherein the first coating structure of the firstblade is selected to provide a first natural vibration frequencydifferent from a second natural vibration frequency of the second blade.11. The method as defined in claim 11, wherein the coating is appliedonly on the first airfoil.
 12. The method as defined in claim 11,wherein the coating is applied on an outer surface of the second airfoilto form a second coating structure, the coating being applied on atleast a portion adjacent the root so that the second blade has a rootcoating thickness and a portion adjacent the tip so that the secondblade has a tip coating thickness, the root coating thickness beinggreater than at least the tip coating thickness and a coating thicknessof the airfoil of the second blade at the mid-span region, wherein thesecond coating structure of the second blade is selected to provide thesecond natural vibration frequency that is different from the firstnatural vibration frequency.
 13. The method as defined in claim 11,wherein the coating is applied so that root coating thickness of thefirst blade is greater than the tip coating thickness of the firstblade.
 14. The method as defined in claim 11, wherein the coating isapplied so that the root coating thickness and the tip coating thicknesson the trailing and the leading edges of the first fan blade is greaterthan the coating thickness of the airfoil at the mid-span region on thetrailing and the leading edges of the first blade.
 15. A compressorrotor for a gas turbine engine, the mistuned compressor rotor comprisingblades circumferentially distributed around and extending a span lengthfrom a central hub, the blades including alternating first and secondblades having corresponding geometric profiles, the first blade havingairfoil with a coating thereon within one or more portions thereof anddefining a first coating structure, the one or more portions of theairfoil including a radially inner portion of the airfoil adjacent ablade root of the first blade and having a root coating thickness and aradially outer portion of the airfoil adjacent a blade tip of the firstblade and having a tip coating thickness, the root coating thicknessbeing greater than at least one of the tip coating thickness and amid-span coating thickness at a mid-span region of the airfoil of thefirst blade, the first coating structure of the first blade selected toprovide a first natural vibration frequency different from a secondnatural vibration frequency of the second blade.
 16. The compressorrotor as defined in claim 15, wherein the airfoil of the second blade isfree of coating.
 17. The compressor rotor as defined in claim 15,wherein the airfoil of the second blade has a coating defining a secondcoating structure, the coating being located on at least a portionadjacent the root of the second blade and having a root coatingthickness and a portion adjacent the tip of the second blade and havinga tip coating thickness, the root coating thickness being greater thanat least the tip coting thickness and a coating thickness of the airfoilof the second blade at the mid-span region, the second coating structureof the second blade selected to provide the second natural vibrationfrequency that is different from the first natural vibration frequency.18. The compressor rotor as defined in claim 15, wherein the rootcoating thickness of the first blade is greater than the tip coatingthickness of the first blade.
 19. The compressor rotor as defined inclaim 15, wherein the root coating thickness and the tip coatingthickness on the trailing and the leading edges of the first blade isgreater than the coating thickness of the airfoil at the mid-span regionon the trailing and the leading edges of the first blade.
 20. A rotorblade for a compressor rotor of a gas turbine engine, the compressorrotor having alternating blades having corresponding geometric profilesbut different coating structures selected to provide different naturalvibration frequencies, the rotor blade comprising an airfoil having ablade root, a blade tip, a coating on one or more portions of theairfoil and defining a first coating structure, the one or more portionsof the airfoil including a radially inner portion adjacent the bladeroot and having a root coating thickness and a radially outer portionadjacent the blade tip and having a tip coating thickness, the rootcoating thickness being greater than at least one of the tip coatingthickness and a mid-span coating thickness defined at a mid-span regionmidway between the blade root and the blade tip.